Space vessels such as spaceships and satellites utilize thrusters to achieve motion in space. A thruster operates on the principle that a force generated in one direction generates an equal force in the opposite direction. By emitting a reaction-mass, a thruster accelerates a spacecraft in the opposite direction. A thruster may be used as a small rocket engine for orbit correction or as the main propulsion of the spacecraft.
Older conventional thrusters used chemical propulsion, which utilized liquid and/or solid propellants. Electric thrusters, which accelerate gases by electrical heating and/or by electric and magnetic field forces, can outperform chemical propulsion systems, in part, because of their high specific impulse (Isp) values. Advantages of electric thrusters include high efficiency and performance, low weight, increased spacecraft orbiting lifetimes, reduced overall costs, and a savings in fuel mass. Advances in onboard electric power sources and smaller more efficient electronic devices have expanded the use of electric thrusters in spacecraft applications.
Electric thrusters that convert electrical energy into kinetic energy may be grouped into three categories: electro thermal propulsion, electrostatic or ion propulsion, and electromagnetic propulsion. Within the electromagnetic propulsion category is the Pulsed Plasma Thruster (PPT), which accelerates the propellant plasma via interaction with an electric arc.
Multiple government and civil entities are developing small and micro sized spacecraft that can benefit from PPTs for space missions. Such spacecraft will require major reductions in thrust levels and/or impulse bits to ensure proper and precise control of the spacecraft. Many missions, in particular those that require significant mission propulsion energies and/or acceleration, will require specific impulses beyond those available from chemical rockets. Because present electric rockets cannot efficiently operate a very low level of power and impulse bits they are not well suited for such missions.
While PPTs are at a high state of development, they generally require high levels of voltage and power to initiate the plasma breakdown and are also very inefficient at low powers when operated at values of expelled propellant velocities of interest to space missions. For example, experimental PPTs have been operated at energy levels down to about 2 joules (J) per pulse requiring the use of high voltage charging supplies which can range from 2,000 to 8,000 volts depending on the design. Also, efficiencies of PPTs decrease with decreasing power and presently, are less than 10 percent efficient when operated at values of propellant velocities of interest to space systems. The inefficiencies result in significant increases in power to achieve desired levels of impulse bits.
An example of such a thruster is shown in FIG. 1 and denoted generally as 10. The thruster 10 fits into the class of propellant devices that operates using an all gas propellent although an all solid solution could also be utilized. In particular, the thruster 10 utilizes a low atomic weight liquid propellant such as water or monopropellant hydrazine (N.sub.2 H.sub.4) or a mixture of two liquids such as water and hydrazine which is stored in the tank 12 and flows through a conduit 14 leading to an opening 16 that forms the feeding mechanism of the thruster 10. The liquid propellent within the tank 12 may be pressurized by high pressure helium in the tank 20, in a manner well known to those of ordinary skill in the art.
The liquid propellent flows through the conduit 14 via the opening 16 and reaches a passage 18 within the thruster 10. The passage 18 leads to a small opening 22 which is sized to provide the correct flow velocity for the liquid propellent and reduce back flow into the passage 18. In the passage 18, the liquid propellent is partially or fully atomized and partially evaporated, so that there is a two phase flow of liquid and gas into the thruster 10. The liquid propellent is disassociated into low atomic weight elemental constituents thereof by an electric discharge that forms a plasma arc within the thruster 10.
The liquid gas and plasma flow from an open end 24 of the passage 18 into the thrust nozzle 30 which, as shown, is shaped as a cone or bell having a curved confining surface, to provide high efficiency and conversion of the high pressure plasma into a directed supersonic flow having high momentum. This discharge of plasma is established primarily by the use of a high voltage DC (HVDC) power supply 32 which is coupled to electrodes 34 and 36 of the thruster 10.
In particular, the thruster 10 operates when liquid from the tank 12 flows into the passage 18 and a high voltage ignition signal supplied by the HVDC power supply 32 is applied at terminals 34 and 36 at a predetermined frequency, such as 200 pulses per second, for example. This ignition voltage can vary but according to one design ranges from 2,000 volts to 8,000 volts. The ignition signal supplied by the HVDC power supply 32 causes a discharge to be established in the passage 18 between the electrodes 34 and 36 at a time when partially atomized fluid is entering the thrust nozzle 30 through the opening 24. The velocity and mass flow rate of liquid flowing through the passage 18 and the repetition rate and energy of the plasma discharge between the electrodes 34 and 36 are matched to achieve optimum operation.
Typically, the HVDC power supply 32 raises the voltage of the thruster 10 until an electrical breakdown occurs between the electrodes 34 and 36. The requirement, however, that the HVDC supply 32 generate high levels of ignition voltages makes the thruster 10 unsuitable for many propulsion applications where small spacecraft are involved. The HVDC supply 32 can be large and not well suited for such applications. Moreover due to its size, the HVDC supply 32 makes it difficult to achieve small and precise maneuvers for some spacecraft missions.
For many space mission applications, where small space systems are involved and which require extremely precise control, the use of high power and/or high voltage ignition circuits is impractical. Examples of such missions are those which require extremely precise ephemeris control and those which are otherwise penalized by high thrust, such as missions which require multiple acceleration and deceleration maneuvers. Thus a PPT that is able to efficiently operate without a high voltage ignitor system and at power levels several orders of magnitude less than prior art designs would be advantageous.